The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor to generate hot combustion gases which flow downstream through one or more turbines which extract energy therefrom. A turbine includes a row of circumferentially spaced apart rotor blades extending radially outwardly from a supporting rotor disk. Each blade typically includes a dovetail which permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk. An airfoil extends radially outwardly from the dovetail.
The airfoil has a generally concave pressure side and generally convex suction side extending axially between corresponding leading and trailing edges and radially between a root and a tip. The blade tip is spaced closely to a radially outer turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades. Maximum efficiency of the engine is obtained by minimizing the tip clearance or gap, but is limited by the differential thermal expansion and contraction between the rotor blades and the turbine shroud for reducing the likelihood of undesirable tip rubs.
Since the turbine blades are bathed in hot combustion gases, they require effective cooling for ensuring a useful life thereof. The blade airfoils are hollow and disposed in flow communication with the compressor for receiving a portion of pressurized air bled therefrom for use in cooling the airfoils. Airfoil cooling is quite sophisticated and may be effected using various forms of internal cooling channels and features, and cooperating cooling holes through the walls of the airfoil for discharging the cooling air.
The airfoil tip is particularly difficult to cool since it is located directly adjacent to the turbine shroud, and the hot combustion gases flow through the tip gap therebetween. A portion of the air channeled inside the airfoil is typically discharged through the tip for cooling thereof. The tip typically includes a radially outwardly projecting edge rib disposed coextensively along the pressure and suction sides between the leading and trailing edges. A tip floor extends between the ribs and encloses the top of the airfoil for containing the cooling air therein.
The tip rib is typically the same thickness as the underlying airfoil sidewalls and provides sacrificial material for withstanding occasional tip rubs with the shroud without damaging the remainder of the tip or plugging the tip holes for ensuring continuity of tip cooling over the life of the blade. The tip ribs, also referred to as squealer tips, are typically solid and provide a relatively large surface area which is heated by the hot combustion gases. Since they extend above the tip floor they experience limited cooling from the air being channeled inside the airfoil. Typically, the tip rib has a large surface area subject to heating from the combustion gases, and a relatively small area for cooling thereof. The blade tip therefore operates at a relatively high temperature and thermal stress, and is typically the life limiting point of the entire airfoil.
Accordingly, it is desired to provide a gas turbine engine turbine blade having improved tip cooling.